This disclosure relates to a gas turbine engine components, such as airfoils. More particularly, the disclosure relates to a cooling arrangement within a cooling cavity of the airfoil, for example.
Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.
Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. The turbine vanes, which generally do not rotate, guide the airflow and prepare it for the next set of blades. Many blades and vanes, blade outer air seals, turbine platforms, and other gas turbine engine components include internal cooling cavities that are supplied cooling fluid to maintain the component within desired operating temperatures.
In order to meet desired aerodynamic performance and structural capability, airfoil shapes in particular have developed into complex three-dimensional geometries where the cross-sectional area varies significantly from root to tip. This cross-sectional variation makes it difficult to maintain high internal heat transfer coefficients throughout the part. Current technology such as microcircuits, which generally produce relatively high flow restrictions, can maintain the high heat transfer coefficients, but require additional flow and cost to do so.